This disclosure relates to a method of modifying the end wall contour in a turbine using laser consolidation. It also relates to the turbine blades derived therefrom.
A turbine system generally comprises stationary vanes and blades rotating about the central axis. Each row of airfoil members divides the annulus of the gas path into a series of sectoral passages, each bounded by the opposed suction and pressure surfaces of an adjacent pair of airfoils and the radially inner and outer walls of the endwall. The inner endwall connected to the rotating blade airfoil root is also termed a platform. The flow field within the sectoral passages is complex and includes a number of secondary vertical flows that act as major sources of energy loss. Reference can be made to Sieverding (1985) “Secondary Flows in Straight and Annular Turbine Cascades”, Thermodynamics and Fluids of Turbomachinery, NATO, Vol. 11, pp 621-624 for a detailed discussion of these flows. The relative importance of these secondary flows increases with the increase of aerodynamic duty or decrease of aspect ratio of the blades. Not only is there energy dissipation in the secondary flows themselves, but they can also affect adversely the fluid flow downstream because they cause deviations to the exit angles of the flow from the rows of airfoil members.
It is found that it is the endwall boundary layers (also known as the end wall contours) that give rise to a substantial part of these secondary flows. FIG. 1 shows a flow model illustration taken from Takeishi et al (1989), “An Experimental Study of the Heat Transfer and Film Cooling on Low Aspect Ratio Turbine Nozzles”, ASME Paper 89-GT-187. This shows part of a row of turbine blades projecting from a cylindrical surface that forms a radially inner endwall of the annular passage from which the blade airfoil extends. The principal flow features as shown in the model in the FIG. 1 are (i) rolling up of the inlet boundary layer L into a horseshoe vortex H at the blade leading edge due to a pressure variation at the intersection of the leading edge and the endwall. The pressure surface side leg of this flow becomes the core of a passage vortex P that is a dominant part of the secondary flow. On the endwall beneath the passage vortex a new boundary layer is formed, indicated as cross-flow B, which starts in the pressure side corner of the endwall of the blade passage. (ii) Upstream of the crossflow B, the inlet boundary layer is deflected across the passage, as indicated by crossflow A. The endwall separation line S marks the furthest penetration of the bottom of the inlet boundary layer A into the blade passage and divides it from the new boundary layer (crossflow B) forming downstream of it. (iii) The new endwall boundary layer, crossflow B, continues onto the blade suction surface until it separates, along an airfoil separation line V, and feeds into the passage vortex P. The horseshoe vortex suction side leg, referred to as the counter vortex U in the FIG. 1, remains above the passage vortex P and moves away from the endwall as the passage vortex grows. (iv) A small counter vortex C may be initiated in the corner region between the blade suction surface and the endwall, rotating in the opposite sense to the passage vortex. (v) Also illustrated in FIG. 1 are the attachment line T which represents the division of the incoming boundary layer flow L between adjacent passages, and the saddle point D, where the attachment line T and the endwall separation line S intersect.
In general, the passage vortex will increase the exit angle of the flow at the endwall (referred to as “over turning”) with the compensatory reduction in exit angle away from the wall (referred to as “under turing”). These effects give rise to deviations of the inlet flow to the next airfoil row, causing the angle of incidence of the flow on the airfoils to vary positively or negatively from the design value and so reduce the aerodynamic efficiency of the flow. They also promote surface heating which is undesirable. Surface heating leads to higher temperatures at the surface.
It is therefore desirable to modify the endwall contour of turbine blades so as to minimize these secondary flows and to improve aerodynamic efficiency as well as to reduce heating to the platform and the blades. A number of publications disclose newer end-wall designs that minimize the formation of these secondary flows to improve the aerodynamic efficiency of the turbine rotor. These newer designs however, require the end-wall modification to be made during the casting process for the blades prior to the assembly of the new blades with the platform. Meeting these design requirements during the casting process is difficult and expensive. For example, the fillet radius between the blade and the platform is generally only sized to meet the minimum requirements set by the casting. Further modifications are often desirable after the casting process in order to accomplish end wall contour modifications.
It is therefore desirable to have a process that facilitates modification of the existing turbine blades that do not contain the aforementioned end-wall modifications. It is also desirable to have a process that facilitates modification of the existing turbine blades that are not hitherto cast to desired specifications. The process can be advantageously used to modify existing blades that are already in service.